Hybrid rocket motor using a turbopump to pressurize a liquid propellant constituent

ABSTRACT

A hybrid rocket motor includes a storage tank which stores an oxidizer under relatively low pressure, a turbopump preferably directly coupled to an outlet of the storage tank which pressurizes the oxidizer to a relatively high pressure, a combustion chamber including a solid fuel, and an injector between the turbopump and combustion chamber through which the oxidizer is injected into the combustion chamber. According to a preferred aspect of the invention, the turbopump is operated by an expander cycle of a heat exchanger. According to another preferred aspect of the invention, the fluid flowing through the heat exchanger is oxidizer tapped from the storage tank. A barrier is maintained between an oxidizer feed line from the turbopump and the injector until sufficient pressure is created by the turbopump to pump the oxidizer at the requisite pressure into the injector.

BACKGROUND OF THE INVENTION

1. Field of the Invention

This invention relates broadly to self-propelled projectiles. Moreparticularly, this invention relates to rockets powered by hybridpropellant systems.

2. State of the Art

Rocket boosters (motors) generally fall into three classes: solidpropellant boosters in which a solid fuel element, or grain, undergoescombustion to produce thrust that propels the rocket, liquid propellantboosters that accomplish the same function with a liquid fuel material,and hybrid boosters, described below. Solid and liquid rocket boosterscan produce relatively large amounts of thrust, but for a relativelyshort amount of time. In addition, solid and liquid rocket boosters aregenerally expensive to develop and produce due to the inherent dangersof the highly combustible solid fuels.

Hybrid rocket boosters are described in detail in co-owned U.S. Pat. No.5,715,675 to Smith et al., which is hereby incorporated by referenceherein in its entirety. They have been characterized as a cross betweena solid propellant booster and a liquid propellant booster. Generallyhybrid boosters use a fluid reactant (an oxidizer) to burn a solid fuelelement, although they may use a combustible liquid fuel and a solidreactant. The hybrid rocket propellant (fuel and reactant together) canbe ignited by an igniter, such as an electrically-generated spark, bypyrotechnic means, or by initial injection of an ignition fluid whichexothermically reacts with the fuel or reactant. The fuel of a hybridrocket is inert until mixed with the oxidizer in the presence of anigniter in a combustion chamber. As such, there is no danger ofinadvertent and uncontrollable combustion.

When the propellant is combusted, the oxidizer must be injected at arelatively high pressure along the surface of the solid reactant toprovide the necessary thrust level. The pressure to inject the oxidizeris created by storing the fluid reactant at a relatively high pressure,e.g., 1000 psi, in a tank. Of course, a tank capable of withstanding1000 psi must have relatively thick walls and is therefore extremelyheavy. The weight of the fluid tank influences rocket flight time anddistance traveled.

SUMMARY OF THE INVENTION

It is therefore an object of the invention to provide a hybrid rocketmotor which uses a storage tank with relatively thinner walls.

It is also an object of the invention to provide a hybrid rocket motorwhich stores fluid reactant at relatively low pressures.

It is another object of the invention to provide a hybrid rocket motorwhich uses a relatively light weight fluid reactant tank.

It is a further object of the invention to provide a hybrid rocket motorwhich injects fluid reactant into a combustion chamber at a relativelyhigh pressure.

In accord with these objects, which will be discussed in detail below, ahybrid motor includes a storage tank which stores fluid reactant(oxidizer) under relatively low pressure, e.g., 100 psi, a turbopumppreferably directly coupled to an outlet of the storage tank whichpressurizes the oxidizer to a relatively high pressure, e.g., 1000 psi,a combustion chamber including a solid fuel, and an injector between theturbopump and combustion chamber through which the oxidizer is injectedinto the combustion chamber. According to a preferred aspect of theinvention, the turbopump is operated by expanded gas from a heatexchanger in an expander cycle. According to another preferred aspect ofthe invention, the fluid flowing through the heat exchanger is the samefluid as the fluid reactant, and more preferably is oxidizer tapped fromthe storage tank. A barrier is maintained between a fluid reactant feedline from the turbopump and the injector until sufficient pressure iscreated by the turbopump to pump the fluid reactant at the requisitepressure into the injector.

A rocket is also provided which incorporates the hybrid motor. Therocket includes a nose cone at the fore end, a rear nozzle, and a casingabout the hybrid motor.

Additional objects and advantages of the invention will become apparentto those skilled in the art upon reference to the detailed descriptiontaken in conjunction with the provided figures.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a broken longitudinal section view of a rocket provided with ahybrid booster according to the invention;

FIG. 2 is a broken schematic of the hybrid motor according to a firstembodiment of the invention;

FIG. 2a is an enlarged broken schematic section of the hybrid motor ofthe first embodiment of the invention;

FIG. 3 is a broken schematic of the hybrid motor according to a secondembodiment of the invention;

FIG. 4 is a broken schematic of the hybrid motor according to a thirdembodiment of the invention;

FIG. 5 is a section view across line 5—5 in FIG. 4; and

FIG. 6 is a broken schematic of the hybrid motor according to a fourthembodiment of the invention.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

Turning now to FIGS. 1 and 2, a rocket 10 includes a hybrid motor 12surrounded by a tubular casing 13, a nose cone 14 at a front end of thecasing, and an exhaust nozzle 16 at an aft end of the casing. The hybridmotor 12 includes a storage tank 20 holding a fluid oxidizer, preferablyliquid oxygen 22, and having an outlet 24, and a pump 28 which operatesto pump the liquid oxygen 22 from the storage tank 20 through a mainline 30 to an injector 32.

A pressurant such as helium or nitrogen 23 is provided in the tank 20 toincrease the tank pressure, e.g., to 100 psi, but does not create thetypical high pressure (e.g., 1000 psi) typically created in storagetanks used in hybrid motors. As such, relatively low structuralrequirements for the tank 20 are necessitated. The motor 12 alsoincludes a combustion chamber 27 provided with a thick-walled tubularcylinder solid fuel grain body 26. The solid fuel grain body 26 ispreferably AP/HTPB (ammonium perchlorate/hydroxyl-terminatedpolybutadiene). The injector 32 preferably extends into the head end 48of the combustion chamber 27 and injects the oxidizer 22 toward thesolid fuel grain body 26.

Referring to FIGS. 2 and 2a, the pump 28 includes a preferably radialimpeller 34 adjacent the outlet 24 which pressurizes the oxidizer fromthe storage tank 20 through volute 36 and into the main line 30. Thecentrifugal impeller 34 is provided at one end of a shaft 35, and apreferably axial turbine 38 is provided at the other end of the shaft.As such, according to a preferred aspect of the invention, the pump 28is preferably a turbopump. The axle 35, impeller 34 and turbine 38 arerotodynamically supported on preferably ceramic/metal bearings 37, 39axially aligned with the outlet 24. A shaft seal 41 is also preferablyprovided between bearing 39 and the turbine 38. The centrifugal impeller34 and the turbine 38 are both axially aligned with the outlet 24 of thetank 20, and the outlet 24 of the tank 20 is preferably directly coupledto the pump 28; i.e., without intervening flexible tubes or othertortuous paths therebetween thereby eliminating the complex array ofplumbing commonly used in rocket motors provided with a turbopump. Suchan arrangement minimizes the pressure drop of fluid flowing from thetank to the pump, which simplifies the design of the pump and improvespump performance.

According to a preferred aspect of the invention, the turbine 38 ispreferably powered in an expander cycle by a heat exchanger 40. The heatexchanger 40 includes an inlet 42 which is in fluid communication withthe main line 30 via a tap 44, but may alternatively be in fluidcommunication with the storage tank 20 from a location above the pump,e.g., from a tap 44 a (shown in broken lines) extending directly fromthe tank 20. A throttle control valve 46 is preferably provided in thetap 44. Alternatively, with the tank tap 44 a, a valve 46 a may also beprovided.

According to a first embodiment of the invention, the heat exchanger 40includes a tubular coil 45 provided around the head end 48 of thecombustion chamber, and preferably the injector 32 is situated to extenda distance into the head end. The heat exchanger outlet 50 feeds into anexpansion chamber 52, and the axial turbine 38 is provided within theexpansion chamber 52. An exit 52 is provided on the other side of theaxial turbine 38 from the expansion chamber 52 for venting expandedoxidizer overboard, and may be used, if desired, for attitude control ofthe rocket. Fluid seals 53 are provided around the tap 40, main line 30,and expansion chamber 52 at the head end of the combustion chamber toprevent liquid oxygen from escaping at the respective locations.

The main line 30 is preferably provided with a flow controller 54 whichobstructs the flow of oxidizer into the injector 32 until it isreoriented, destroyed or otherwise at least partially removed from themain line path. The flow controller 54 may be a valve, a burst discdestroyed by pressure, pyrotechnics or otherwise, or the flowcontrollers (barriers and valves) disclosed in U.S. Pat. Nos. 6,058,697and No. 6,125,763, which are hereby incorporated by reference herein inits entirety. In operation, with the flow controller 54 and tap valve 46in closed configurations, liquid oxygen 22 flows from the storage tank,around the centrifugal pump 34, and into the main line 30 when it isstopped by the flow controller 54. When the tap valve 46 is then opened,the liquid oxygen enters the inlet 42 of the heat exchanger 40 and flowsinto the tubular coils 45 of the heat exchanger. As the liquid oxygen 22is a cryogenic fluid and the exterior of the heat exchanger is initiallyat ambient temperature, there is a temperature differential of hundredsof degrees therebetween which causes the liquid oxygen to rapidly warm.The liquid oxygen then exits the outlet 50 of the heat exchanger andenters the expansion chamber 52, where the liquid oxygen (LOX) undergoesa flash expansion to gaseous oxygen (GOX). This energy of the flashexpansion rotates the axial impeller 38 which rotates the centrifugalpump 34. As the impeller 38 continues to rotate, the pump 34 increasesthe pressure of the liquid oxygen on the main line side of the flowcontroller 54. Once sufficient pressure, e.g., 1000 psi, is created toprovide a hybrid motor with suitable thrust, the flow controller 54 isreconfigured, destroyed or otherwise at least partially removed topermit the oxidizer to flow to the injector 32. The injector 32 theninjects the highly pressurized liquid oxygen into the combustion chamber27. When the oxidizer 22 is combusted with the solid fuel grain 26, therocket is provided with thrust.

It should be appreciated that with the provided arrangement the bearings37, 39 do not require gas-tight seals. This is in contrast to prior arthybrid rocket motors such as disclosed in U.S. Pat. No. 5,572,864 toJones which include a turbopump powered by steam generated by a separatemotor. In the prior art, different fluids are located on either side ofthe turbine and the pump (steam and liquid oxygen), and must not be incommunication for proper operation of the rocket motor. This necessarilyrequires the use of gas-tight seals at the bearings between the turbineand the centrifugal pump. However, such seals are difficult to maintainas they must operate across large temperature differentials: theoxidizer is preferably a cryogenic fluid, while the turbine operatingfluid, e.g., steam, is much warmer. Seals which operate across suchconditions are very expensive and prone to leakage or failure. Thepresent invention does not require the use of gas-tight seals, as oxygen(GOX and LOX) is located on both sides of the turbopump.

Turning now to FIG. 3, a second embodiment of a turbopump 128 operatedin an expander cycle by a heat exchanger and for use in a hybrid rocketmotor is shown. The heat exchanger 140 includes two preferablyconcentric tubes 145, 146. The inner tube 145 includes an open end 160,and the outer tube 146 includes a closed end 162. The open end 160 isprovided adjacent the closed end 162. The liquid oxygen is fed into theinner tube 145 and out the open end 160, and then travels up the annularspace 164 between the inner and outer tubes to an expansion chamber 152housing the axial turbine 138. A valve 144 is provided to control theflow of oxygen to the tubes 145, 146.

Turning now to FIG. 4, a third embodiment of a turbopump 228 operated inan expander cycle by a heat exchanger and for use in a hybrid rocketmotor is shown. Rather than using a tubular coil around the periphery ofthe head end of the combustion chamber, as described with respect to thefirst embodiment, the heat exchanger is integrated into the injector232. Referring to FIGS. 4 and 5, the injector 232 includes a faceportion 260 defining a circuitous path 245, and injector holes 264extending through face portion 260, but not intersecting the path 245.The injector 232 also includes an inlet 242 and an outlet 250communicating with the path 245.

The inlet 242 is coupled to a tap 244 which receives liquid oxygen. Theliquid oxygen flows from the tap 244 to the inlet 242, and through thepath 245 to the outlet 250 where it is then expanded in an expansionchamber 252 and causes rotation of the axial impeller 238.

Turning now to FIG. 6, a fourth embodiment of a turbopump 328 operatedin an expander cycle by a heat exchanger and for use in a hybrid rocketmotor is shown. As in the third embodiment, the heat exchanger 340 isintegrated into the injector 332. More particularly, the injector 332extends into the head end of the combustion chamber 348 and includes aconcentric arrangement of an aft portion of an inner inlet 370 and anaft portion of an outer outlet 372. A forward portion of the inlet 370extends through a portion of the outlet and is sealed in communicationwith the main line 330. A valve 376 is provided in either the main line330 or the inlet 370, or at a juncture of the two. The face 360 of theinjector 332, provided with a plurality of holes 364, is located at anend of the inlet 370, and a burst disc 354 or other removable barrier ispreferably provided over the holes 364 on the face 360 of the injector.The outlet 372 extends upwards and expands to form a fluid expansionchamber 352. A plurality of preferably radially oriented taps 378 placethe inner and outer pathways 370, 372 in fluid communication adjacentthe face 360 of the injector.

When valve 376 is opened, liquid oxygen flows from the tank 320 into themain line 330 and then into the inlet 370 of the injector 332. Theoxygen flows through the taps 378 in the wall of the injector and intothe outlet 372, where the oxygen is quickly heated and expanded intogaseous oxygen. When the gaseous oxygen enters the expansion chamber352, it rapidly expands and results in rotation of the axial impeller338, which thereby operates the pump 328. Once sufficient pressure iscreated in the inner pathway 370 at the injector face 360 by the pump328, the barrier 354 is removed, e.g., by bursting at a desiredpressure, such that the oxygen is injected through the holes 364 in theface 360 of the injector and into the combustion chamber 348. Thecontinual feed of oxygen from the inlet 370 into the taps 378 and up theoutlet 372 continually operates the pump 328 and maintains the injectedoxygen at a highly pressurized state.

There have been described and illustrated herein embodiments of a hybridrocket booster and a rocket provided with the booster. While particularembodiments of the invention have been described, it is not intendedthat the invention be limited thereto, as it is intended that theinvention be as broad in scope as the art will allow and that thespecification be read likewise. Thus, while the preferred oxidizer isliquid oxygen, it will be appreciated that other non-self pressurizingoxidants such as red fuming nitric acid (RFNA), nitrogen tetroxide(NTO), and hydrogen peroxide (H₂O₂) may also be used. While the hybridfuel grain is preferably HTPB, other fuel grains known in the art, suchas ABS resin, CTPB, PBAN or other fuel/binder systems. In addition,while in the first embodiment a tubular coil of the heat exchanger isprovided around the periphery of the head end of the combustion changer,it may be otherwise located, e.g., about the injector, or spaced-apartfrom both the periphery and the injector. Also, the heat exchanger neednot be coiled, but may be provided in another circuitous path adjacentor within the combustion chamber. Furthermore, while the turbine isshown and described in an axial configuration, it will be appreciatedthat a radial inflow turbine may be used instead. It will therefore beappreciated by those skilled in the art that yet other modificationscould be made to the provided invention without deviating from itsspirit and scope as so claimed.

What is claimed is:
 1. A hybrid rocket motor, comprising: a) a containerhaving a fluid reactant therein and an outlet; b) a combustion chambercontaining a solid reactant therein; c) an injector between saidcontainer and said combustion chamber; and d) a turbopump including aturbine and a pump axially aligned with said outlet, said pump adaptedto increase a pressure of said fluid reactant exiting said outlet andinjected through said injector into said combustion chamber.
 2. A hybridrocket motor according to claim 1, wherein: said pump is directlyconnected to said outlet.
 3. A hybrid rocket motor according to claim 1,further comprising: bearings which support said turbine and said pump,wherein said bearings are not provided with gas-tight seals.
 4. A hybridrocket motor according to claim 1, further comprising: e) a heatexchanger at least partially provided in said combustion chamber, saidheat exchanger having an inlet and an outlet; and f) an expansionchamber coupled to said outlet of said heat exchanger, said turbinebeing provided in said expansion chamber.
 5. A hybrid rocket motoraccording to claim 1, further comprising: d) an at least partiallyremovable barrier between said pump and said injector.
 6. A hybridrocket motor, comprising: a) a container having a fluid reactant thereinand an outlet; b) a combustion chamber containing a solid reactanttherein; c) an injector between said container and said combustionchamber; and d) a heat exchanger at least partially within saidcombustion chamber.
 7. A hybrid rocket motor according to claim 6,further comprising: e) a turbopump including a turbine and a pump, saidpump adapted to increase a pressure of fluid reactant exiting saidoutlet and injected through said injector into said combustion chamber.8. A hybrid rocket motor according to claim 7, wherein: said pump isdirectly connected to said outlet.
 9. A hybrid rocket motor according toclaim 7, wherein: said turbopump includes bearings which support saidturbine and said pump, and wherein said bearings are not provided withgas-tight seals.
 10. A hybrid rocket motor according to claim 7, furthercomprising: f) an at least partially removable barrier between said pumpand said injector.
 11. A projectile, comprising: a) a motor having aforward end and an aft end, said motor including, i) a container havinga fluid reactant therein and an outlet, ii) a combustion chambercontaining a solid reactant therein, iii) an injector between saidcontainer and said combustion chamber, and iv) a turbopump including aturbine and a pump axially aligned with said outlet, said pump adaptedto increase a pressure of said fluid reactant exiting said outlet andinjected through said injector into said combustion chamber; b) atubular casing around said motor; c) a nose portion coupled to saidforward end of said motor; and d) a nozzle coupled to said aft end ofsaid motor.
 12. A projectile, comprising: a) a motor having a forwardend and an aft end, said motor including, i) a container having a fluidreactant therein and an outlet, ii) a combustion chamber containing asolid reactant therein, iii) an injector between said container and saidcombustion chamber, and iv) a heat exchanger at least partially withinsaid combustion chamber; b) a tubular casing around said motor; c) anose portion coupled to said forward end of said motor; and d) a nozzlecoupled to said aft end of said motor.
 13. In a hybrid rocket motorhaving a container storing a fluid oxidizer and having a first outlet, acombustion chamber containing a solid fuel grain therein, a turbopumpincluding a turbine side and a pump side, said pump side adapted toincrease a pressure of the fluid oxidizer exiting said first outlet andinjected into said combustion chamber, the improvement comprising: agaseous form of the oxidizer on said turbine side of said turbopump, anda liquid form of said oxidizer on said pump side of said turbopump. 14.The improvement of claim 13, wherein: said oxidizer is oxygen.
 15. Theimprovement of claim 13, wherein: said liquid oxidizer is converted intosaid gaseous oxidizer using an expansion cycle of a heat exchanger.